Aircraft rotor blade of shape adapted for acoustic improvement during an approach flight and for improving performance in forward flight

ABSTRACT

A blade of a rotor for a rotary wing aircraft. The blade presents relationships for variation in the sweep and in the chord of the profiles of sections of the blade, in particular in order to improve the twisting stiffness and the bending stiffness of the blade. The blade is then double-tapered and it presents three sweeps, making it possible firstly to improve the aerodynamic performance of the blade in forward flight, and secondly to reduce the noise given off by the blade, in particular during an approach flight.

CROSS REFERENCE TO RELATED APPLICATION

This application claims priority to French patent application No. FR 1502662 filed on Dec. 21, 2015, the disclosure of which is incorporated inits entirety by reference herein.

BACKGROUND OF THE INVENTION

(1) Field of the Invention

The present invention relates to the field of lift-generating airfoilsurfaces and more particularly to airfoil surfaces forming a rotarywing.

The present invention relates to a blade for a rotary wing aircraftrotor and to a rotor having at least two such blades. The blade isintended more particularly for a main rotor for providing a rotary wingaircraft with lift and possibly also with propulsion.

(2) Description of Related Art

Conventionally, a blade extends longitudinally along its span from afirst end for fastening to a rotary hub of a rotor to a second endreferred to as its “free” end. Relative to the rotor, it can beunderstood that the blade extends radially from the first end towardsthe second end in a spanwise direction. Furthermore, the blade extendstransversely from a leading edge towards a trailing edge of the bladealong the chord of the blade.

The blade is thus driven in rotation by a rotary hub of the rotor. Theaxis of rotation of the hub thus corresponds to the axis of rotation ofthe blade.

The first end is generally referred to below by the term “blade start”,while the free second end is referred to by the term “blade tip”.

In operation, each blade of a rotor is subjected to aerodynamic forces,in particular an aerodynamic lift force during the rotary motion of therotor, which force serves to provide the aircraft with lift, or indeedwith propulsion.

For this purpose, the blade has an airfoil portion situated between theblade start and the blade tip. This airfoil portion is constituted by asuccession of airfoil profiles along the span direction, which airfoilprofiles are often referred to below for short as “profiles”. Eachprofile is situated in a transverse plane that is generallyperpendicular to the span direction and it defines a section of theblade. This airfoil portion is arranged between the start of the airfoilportion and the blade tip and it provides substantially all of the liftof the blade.

The shape of the transition zone between the blade start and the startof the airfoil portion is generally imposed by manufacturing constraintsand by structural constraints concerning the blade. This transition zonebetween the blade start and the start of the airfoil portion may bereferred to by the term “blade root” and its aerodynamic performance ismuch less than the aerodynamic performance of the airfoil portion. Thestart of the airfoil portion is thus situated between the blade startand the blade tip, in the vicinity of the blade root. This transitionzone may nevertheless generate some lift force. In addition, thistransition zone, which is situated in the vicinity of the hub of therotor, nevertheless provides some small contribution to the total liftof the blade, regardless of its aerodynamic shape.

For example, the profiles of the sections of the blade in the airfoilportion are characterized by a thin trailing edge, ideally of zerothickness, whereas the trailing edge in the vicinity of the blade startand of the transition zone between the blade start and the start of theairfoil portion is thick, and possibly rounded.

A rotary wing aircraft presents the advantage of being capable of flyingboth with high forward speeds during cruising flight and also withforward speeds that are very low, and it is also capable of performinghovering flight. A rotary wing aircraft thus presents the advantage ofbeing able to land on zones of small area and thus, for example, closerto inhabited areas or indeed on landing decks or pads.

Nevertheless, forward flight at high speeds requires the blades to haveaerodynamic characteristics that may be different from, or evenunfavorable for the characteristics needed for flight at very lowforward speeds and for hovering flight.

Likewise, the aerodynamic characteristics of blades also influence thenoise generated by blades. Such noise can be problematic during stagesof approaching and landing because of the proximity of inhabited areas.Furthermore, strict acoustic certification standards lay down soundlevels with which rotary wing aircraft are required to comply.

For a predetermined selection of airfoil profiles, the geometricalcharacteristics of a blade that have an influence on the aerodynamicperformance of the blade during forward flight at high speeds and duringhovering flight and also on the acoustic signature of the blade areconstituted in particular by the chords of the airfoil profiles of thesections of the blade, by the sweep of the blade, and by the twist ofthe blade.

It should be recalled that the chord is the distance between the leadingedge and the trailing edge of the profile of a blade section. This chordmay vary along the span of the blade. The term “taper” is generally usedto designate a reduction in chords going along the span of the blade,however this term can also designate an increase in chords along thespan of the blade.

Sweep may be defined as the angle formed by the leading edge of theblade with a particular axis of the blade. By convention, in a zone withforward sweep the leading edge forms a sweep angle relative to saidblade axis that is positive in the direction of rotation of the rotor,whereas in a zone with backward sweep the leading edge forms a sweepangle relative to said blade axis that is negative. Said blade axisgenerally coincides with the pitch or feathering axis of the blade.

The twist of a blade consists in varying the setting of the profiles ofthe sections of the blade along the span of the blade. The term“setting” designates the angle formed between the chord of each profileof the sections of the blade relative to a reference plane of the blade,and this angle is referred to as the “twist” angle. By way of example,the reference plane may be the plane perpendicular to the axis ofrotation of the blade and including said blade axis.

The term “twist relationship” designates how the twist angle variesalong the span of the blade. In conventional manner, twist is measuredas being negative when the leading edge of a section profile of theblade is lower than said reference plane.

Effective solutions are known for independently improving theperformance of a blade for high-speed forward flight and the performanceof a blade for hovering flight, as well as the acoustic performance ofthe blade during approach stages.

For example, improving the aerodynamic performance of a blade forhovering flight is characterized by reducing the power drawn by theblade for unchanging rotor lift. This improvement can be obtained bypassive changes to the shape of the blade, in particular by increasingits twist.

An appropriate increase in the twist of the blade enables lift to bedistributed more uniformly over the entire surface area of the blade andconsequently of the rotor, thereby making it possible to reduce thepower absorbed by each blade of the rotor in hovering flight. It shouldbe recalled that increasing twist consists in lowering the leading edgerelative to said reference plane and doing so more towards the blade tipthan towards the blade start because of the variation in thecircumferential speed of the air stream as a function of span. Theaerodynamic performance of the blade in hovering flight is increased inparticular by making the speeds induced along the span of the blade moreuniform in this way.

Nevertheless, when the rotary wing aircraft is traveling at high speed,a large amount of blade twist can lead to the blade tip having negativelift, i.e. generating a lift force that is in the same direction asgravity, for a blade that is in an azimuth position known to the personskilled in the art as an “advancing” blade. The aerodynamic performanceof the blade is thus degraded in forward flight. Furthermore, the levelsof the aerodynamic loads to which the blade is subjected and also thelevels of vibration are likewise greatly increased during forwardflight.

Adding a dihedral at the blade tip also serves to improve theaerodynamic performance of the blade in hovering flight. A dihedral isformed by a blade surface at the blade tip that slopes upwards ordownwards. In hovering flight, the dihedral serves to ensure that thetip vortex generated by any one blade has reduced influence on thefollowing blades of the rotor. Nevertheless, such a dihedral may giverise to a reduction in the aerodynamic performance of the blade inforward flight and also to an increase in vibration.

Furthermore, improving the aerodynamic performance of a blade in forwardflight is characterized by reducing the power consumed by each blade ofthe rotor for given lift and forward speed. This improvement may beobtained by passive modifications to the shape of the blade, and inparticular by modifying its chord along the span of the blade and/or bydecreasing its twist.

For example, the chord of profiles of sections of the blade increasesgoing from the blade start along the span, and then it decreases beforereaching the blade tip. The blade is said to be a “double-tapered”blade. Document EP 0 842 846 describes a double-tapered blade in whichthe maximum chord is situated at a distance lying in the range 60% to90% of the total span of the blade from the axis of rotation of theblade.

Nevertheless, the use of a double-tapered blade often gives rise to anincrease in noise during approach flight as a result in the increasingintensity of vortices given off by and then impacting against eachblade. The use of such a double taper also gives rise to degradedperformance in hovering flight compared with a blade having the sametwist and the same “blade solidity”, which term designates the ratio ofthe total area occupied by the blades of the rotor seen from above tothe area of the rotor disk, i.e. the area that is swept by a blade ofthe rotor on rotating through one revolution.

Furthermore, and in compliance with the above, a decrease in the twistof the blade leads to an increase in the aerodynamic angles of attack atthe blade tip on the advancing blade side. The angles of attack for anon-twisted blade tip are thus closer to zero on the advancing bladeside, thus serving firstly to reduce the negative lift at the blade tipon the advancing blade side and also to reduce local drag, in particularthe drag associated with the appearance of shock waves.

In contrast, reducing twist at the end of the blade leads to a reductionin the stall margin of the blade on the retreating blade side. Inaddition, this reduction in the twist of the blade is unfavorable inhovering flight, as mentioned above.

Documents U.S. Pat. No. 7,252,479 and EP 0 565 413 describe a bladeadapted to high-speed forward flight, combining a double-tapered bladewith a twist relationship.

Finally, the improvement in acoustic performance of a blade duringapproach flight may be characterized by reducing the noise that isgenerated by the interaction between the blade and the air vortexgenerated by the preceding blades of the rotor. This improvement may beobtained by passive modifications to the shape of the blade, inparticular by modifying its sweep along its span.

By way of example, as described in Documents EP 1 557 354, US2012/0251326, and U.S. Pat. No. 6,116,857, a blade with a first zonethat is forwardly-swept and a second zone that is backwardly-sweptavoids the leading edge of a blade in these first and second zones beingparallel to the lines of vortices given off by the preceding blades.Such a blade can thus limit interactions between the blade and thesevortices, e.g. reducing the intensity of impulse noise associated withthe interaction between the blade and the vortices, and consequentlylimiting the appearance of noise.

Furthermore, that blade with two sweeps may also include taper in thebackwardly-swept second zone that also serves to reduce the noise levelgenerated in flight. Specifically, for a given profile, the thickness ofthe blade decreases with shortening chord, thereby decreasing theappearance of so-called “thickness” noise. Likewise, since the area ofthe blade is reduced as a result of its taper, its lift is alsomodified, which can reduce the appearance of so-called “load” noise.

It is also possible to act on the aerodynamic load at the blade tip inorder to modify the vortices given off in the wake of the blade, andconsequently reduce the sound level of the blade. For this purpose, therelationships for variation in the twist and in the chord of the secondprofiles of the blades are modified. Nevertheless, such variations areincompatible with the above-mentioned optimizations concerning hoveringflight or forward flight.

Furthermore, independently of the shape of the blade, it is alsopossible to modify the speed of rotation of the blade or indeed to adoptspecific approach flight paths for the aircraft referred to as “leastnoise approach flight paths” in order to reduce the noise radiated tothe ground by the blades of the aircraft.

Nevertheless, modifying the speed of rotation of the blade makes thework of dynamically balancing the blade more complex. Furthermore, areduction in the speed of rotation of the blade can give rise inparticular to an increase in aerodynamic stalls at the ends of theblade, and consequently to an increase in the dynamic control forces ofthe blade.

It is also possible to combine applying two sweeps with variations inthe chord of the profiles of the sections of the blade with a twistrelationship that is adapted either to hovering flight or else toforward flight. Thus, documents EP 1 557 354 and US 2012/0251326describe blades that are adapted for hovering flight while also enablinga reduction in the noise generated during approach flights. Likewise,document EP 0 842 846 describes a blade that is adapted for forwardflight at high speeds and that enables noise to be limited duringapproach flight.

Nevertheless, the aerodynamic performance of such blades is notoptimized for the stage of flight for which the blades are not adapted.Significantly reducing the noise given off by a blade is in any eventgiven precedence, and the aerodynamic performance of the blade may bedegraded during certain stages of flight. This degradation is due inparticular to a lack of twisting stiffness and/or of bending stiffnessof the blade which can then deform under the aerodynamic and inertialforces to which it is subjected.

In contrast, optimizing blade profiles for high-speed forward flight isdifferent and appears to go against optimizing those profiles forhovering flight. Optimizing profiles both for hovering flight and forhigh-speed forward flight is particularly complex to achieve, since theaerodynamic conditions encountered by the blade are different.Furthermore, during rotation of the rotor, the position of a bladealternates between advancing and retreating relative to the air stream,thereby increasing the differences between the aerodynamic conditionsencountered by the blade.

Finally, the document entitled “Multiobjective-multipoint rotor bladeoptimization in forward flight conditions using surrogate-assistedmemetic algorithms”, given to the “European Rotorcraft Forum” atGallarate (Italy) in September 2011 compares several methods ofoptimizing a blade in forward flight. The blade may have only a twistrelationship, or it may present a combination of relationships forvarying chord and sweep, or indeed it may present a combination ofrelationships for varying twist, chord, and sweep.

BRIEF SUMMARY OF THE INVENTION

An object of the present invention is to overcome the above-mentionedlimitations and to propose a blade that improves the aerodynamicperformance of the blade while reducing the noise given off by the bladeduring an approach flight. The invention also relates to a rotor for arotary wing aircraft having at least two such blades.

The present invention thus provides a blade for a rotor of a rotary wingaircraft, the blade being for rotating about an axis of rotation A, theblade extending firstly along a blade axis B between a blade startsuitable for being connected to a hub of the rotor and a blade tipsituated at a free end of the blade, and secondly along a transverseaxis T substantially perpendicular to the blade axis B between a leadingedge and a trailing edge, the blade comprising an airfoil portionsituated between the blade start and the blade tip, the airfoil portionbeing constituted by a succession of airfoil profiles, each airfoilprofile being situated in a transverse plane substantially perpendicularto the blade axis B and defining a section of the blade, the blade tipbeing situated at a reference distance equal to a rotor radius R fromthe axis of rotation A, a maximum distance between the leading edge andthe trailing edge in the transverse plane constituting a chord c foreach airfoil profile of the blade, a mean chord c being a mean value ofthe chord c over the airfoil portion, a forward first direction beingdefined from the trailing edge to the leading edge, and a rearwardsecond direction being defined from the leading edge to the trailingedge.

The mean chord c is preferably defined by a radius squared r² weightingof the profile of each of the sections of the blade in application ofthe formula:

$\overset{\_}{c} = \frac{\int_{R_{0}}^{R}{{L(r)} \cdot r^{2} \cdot {dr}}}{\int_{R_{0}}^{R}{r^{2} \cdot {dr}}}$

where L(r) is the length of the local chord of a profile of the bladesituated at a radius r from the axis of rotation A, R₀ being the radiusof the start of the airfoil portion, and R being the radius of the bladetip.

Nevertheless, the mean chord c may be defined by an arithmetic mean ofthe chord c of the sections of the blade over all of the airfoil portionof the blade.

This blade of the invention is remarkable in that it presents acombination of relationships for variation in its chord and in itssweep, its chord increasing between the start of the airfoil portion anda first section S1 situated at a first distance from the axis ofrotation A lying in the range 0.6R to 0.9R, its chord decreasing beyondthe first section S1, the sweep of the blade being initially directedtowards the front of the blade between the start of the airfoil portionand a second section S2 situated at a second distance from the axis ofrotation A lying in the range 0.5R to 0.8R, the leading edge forming aforward first sweep angle lying in the range 0° to 10° relative to theblade axis B, the sweep then being directed towards the front of theblade between the second section S2 and the third section S3 situated ata third distance from the axis of rotation A lying in the range 0.6R to0.95R, the leading edge forming a forward second sweep angle lying inthe range 1° to 15° relative to the blade axis B, and the sweep finallybeing directed towards the rear of the blade between the third sectionS3 and the blade tip, the leading edge forming a backward third sweepangle lying in the range −35° to −15° relative to the blade axis B.

This blade of the invention is preferably for the main rotor providing arotary wing aircraft with lift and possibly also with propulsion. Theaxis of rotation A of the blade corresponds to the axis of rotation ofthe hub of the rotor.

The airfoil portion of the blade provides the main part of the lift fromthe blade while the blade is rotating about the axis A. The start ofthis airfoil portion is characterized in particular by a thin trailingedge, whereas between the blade start and the start of the airfoilportion, the trailing edge is thick, or even rounded. The start of thisairfoil portion is thus generally distinct from the blade start and issituated between the blade start and the blade tip, in the vicinity ofthe blade start.

The blade tip is situated at a reference distance equal to the rotorradius R from the axis of rotation A, and this rotor radius R is usedfor locating a profile or indeed a section of the blade along the bladeaxis B. For example, the blade start is situated at a fourth distancelying in the range 0.05R to 0.3R from the axis of rotation A and thestart of the airfoil portion of the blade is situated at a fifthdistance lying in the range 0.1R to 0.4R from the axis of rotation A.The fifth distance is greater than or equal to the fourth distance.

Likewise, the mean chord c of the blade over the airfoil portion is usedto define the chord of each profile of the blade along its span.

The relationship for variation in sweep thus defines a blade with threesweeps that serve advantageously to improve the acoustic signature ofthe blade. The forward first sweep angle is preferably strictly greaterthan 0°.

These three sweeps advantageously prevent the leading edge of a bladebeing parallel with the vortices given off by the preceding blade duringrotation of a blade. These three sweeps thus enable the intensity of theacoustic energy that is generated by interaction between a blade and theair vortices given off by the preceding blades of the rotor to bedecreased over a portion of the span of the blade, in particular duringan approach flight.

In addition, the ends of the preceding blades give off vortices thatform vortex lines of helical shape. It is then advantageous to limit thespan portions of the leading edge of the blade that interactsimultaneously with these vortex lines in order to limit the effect onthe human ear of the noise generated.

Specifically, with a leading edge of the blade having a sweep angle thatvaries continuously over one or more zones defined by the second andthird sections S2 and S3 and also by the blade tip, interaction betweenthe leading edge and the vortices given off by the blades preceding afollowing blade take place simultaneously at a plurality of points onthe leading edge giving rise to the appearance of acoustic energy. Thisresults in impulse noise being given off that is troublesome for thehuman ear, with this phenomenon being penalizing for acousticcertification.

Advantageously, with a straight leading edge of the blade that slopesrelative to the blade axis over each zone defined by the second andthird sections S2 and S3 and by the blade tip, the interaction betweenthe leading edge and these vortices takes place simultaneously at asmaller number of points along the leading edge. This results in adecrease in the impulse nature of the signal given off, making it lesstroublesome for the human ear.

Consequently, the leading edge of the blade is preferably straight andinclined over each zone defined by the second and third sections S2 andS3 and by the blade tip so as to reduce the acoustic energy perceived byan observer.

The sweep is thus preferably formed by a forward first sweep angle, aforward second sweep angle, and a backward third sweep angle, all threeof which are constant respectively between the start of the airfoilportion and the second section S2, then between the second section S2,and the third section S3, and finally between the third section S3 andthe blade tip.

Likewise, the forward first sweep angle α₁ is preferably different fromthe forward second sweep angle α₂ so as to guarantee that there arethree distinct sweeps present along the blade of the invention.

In addition, the forward first sweep angle α₁ may be strictly less thanthe forward second sweep angle α₂ in order to guarantee progressivenessover the two forward sweep zones.

For example, the forward first sweep angle may be equal to 4°, theforward second sweep angle may be equal to 8°, and the backward thirdsweep angle may be equal to −23°.

In the relationship for variation in the chord of the profiles of thesections of the blade, the chord varies about the mean chord c by ±40%between the start of the airfoil portion and the first section S1. Thechord thus varies over the range 0.6c to 1.4c respectively from thestart of the airfoil portion to the first section S1. The variation inthe chord may also be smaller between the start of the profile portionand the first section S1, in particular to penalize the aerodynamicperformance of the blade during hovering flight to a smaller extent. Byway of example, the chord may vary by ±20% about the mean chord cbetween the start of the airfoil portion and the first section S1.

In addition, the chords of the section profiles of the blade arepreferably smaller than the mean chord c over a first portion of theblade, e.g. between the start of the airfoil portion of the blade and afourth section S4 situated at an sixth distance from the axis ofrotation A lying in the range 0.5R to 0.8R. The chords of the sectionprofiles of the blade are then greater than the mean chord c between thefourth section S4 and a fifth section S5 situated at a seventh distancefrom the axis of rotation A lying in the range 0.85R to 0.95R, and thenless than the mean chord c beyond the fifth section S5 to the blade tip.By way of example, the chord of the section profile of the blade in thevicinity of the start of the airfoil portion of the blade lies in therange 0.4 c to 0.9c, while the chord of the section profile of the bladeat the blade tip may lie in the range 0.2c to 0.8c.

Furthermore, the chord may decrease in non-linear manner beyond a sixthsection S6 to the blade tip, this sixth section S6 being situated at aneighth distance from the axis of rotation A lying in the range 0.9R to0.95R. Preferably, the chords of the profiles of the sections of theblade decrease along a curve that is parabolic beyond the sixth sectionS6. It is then possible to refer to a “parabolic” tip cap being presentat the blade tip. Other non-linear shapes are also possible for thisblade tip using polynomial curves such as a Bezier curve.

Under such circumstances, the chord of the profile of the section at theblade tip lies in the range 0.2c₁ to 0.8c₁ where c₁ is the value of thechord of the profile of the section of the blade at the sixth sectionS6, i.e. at the beginning of this zone of non-linear reduction in thechord of the profiles of the sections of the blade. The chord at theblade tip is preferably equal to 0.3c₁.

Combining relationships for variation in the sweep and in the chord ofthe section profiles of the blade serves to improve the aerodynamicperformance in forward flight while reducing the noise given off by theblade, in particular during approach flights.

Furthermore, the blade may combine a twist relationship with therelationships for variation in its chord and in its sweep. In the twistrelationship, the twist of the profiles of the sections of the bladedecreases between a seventh section S7 situated between a ninth distancefrom the axis of rotation A lying in the range 0.3R to 0.4R and theblade tip, a first twist gradient lying in the range −25°/R to −4°/Rbetween the seventh section S7 and an eighth section S8 situated at atenth distance from the axis of rotation A lying in the range 0.4R to0.6R, a second twist gradient lying in the range −25°/R to −4°/R betweenthe eighth section S8 and a ninth section S9 situated at an eleventhdistance between the axis of rotation A lying in the range 0.65R to0.85R, a third twist gradient lying in the range −16°/R to −4°/R betweenthe ninth section S9 and the tenth section S10 situated at a twelfthdistance from the axis of rotation A lying in the range 0.85R to 0.95R,and a fourth twist gradient lying in the range −16°/R to 0°/R betweenthe tenth section S10 and the blade tip.

Advantageously, combining the relationships for variation in the chordand the sweep of the profiles of the sections of the blade with therelationship for twist serves to improve the aerodynamic performance ofthe blade mainly in hovering flight without degrading either theaerodynamic performance of the blade in forward flight or the noisegenerated by the blade during approach flights. Specifically, twist isimportant in a first zone of the blade, e.g. in the range 0.3R to 0.7R,and thus serves to compensate for the small chord that is essentiallyless than the mean chord c. Furthermore, the untwisting in a second zoneof the blade, e.g. in the range 0.7R to 0.9R, is favorable to forwardflight for an advancing blade, but gives rise to an increase in forceson a retreating blade. Advantageously, in this second zone, the chordsof the profiles of the sections of the blade are essentially greaterthan the mean chord c, thus making it possible to accommodate theseincreased forces without degrading the aerodynamic behavior of theretreating blade.

The relationship for variation in the twist of the blade may bepiecewise linear, i.e. between adjacent pairs of sections selected fromthe sections S7, S8, S9, and S10, and between the tenth section S10 andthe blade tip, or it may be non-linear over all of the airfoil portionof the blade.

When the twisting relationship is piecewise linear, this twistingrelationship is constituted by straight line segments, a segmentcharacterizing variation of twist between two adjacent segments fromamong the sections S7, S8, S9, and S10, and between the tenth sectionS10 and the blade tip. The twist gradient, which is the local derivativeof the twist along the span of the blade, then corresponds to the slopesof the straight lines supporting these segments. This twist gradient isthen formed by discontinuous horizontal lines, a line being situatedbetween adjacent sections and between the tenth section S10 and theblade tip.

Furthermore, in order to enable twist variation to be compatible bothwith hovering flight and with forward flight and also with therelationship for variation in chord, the first twist gradient situatedbetween the seventh section S7 and the eighth section S8 is preferablyless than the second twist gradient situated between the eighth sectionS8 and the ninth section S9, the second twist gradient is preferablygreater than the third twist gradient situated between the ninth sectionS9 and the tenth section S10, and the third twist gradient is preferablyless than the fourth twist gradient situated between the tenth sectionS10 and the blade tip.

When the twist relationship is non-linear over the airfoil portion, thetwist gradient is preferably a curve that is continuous over the entireairfoil portion of the blade. The first twist gradient then reaches afirst plateau lying in the range −25°/R to −15°/R in the vicinity of theeighth section S8, the second twist gradient reaches a second plateaulying in the range −14°/R to −4°/R in the vicinity of the ninth sectionS9, the third gradient of the twist reaches a third plateau lying in therange −16°/R to −6°/R in the vicinity of the tenth section S10, and thefourth gradient of the twist lies in the range −10°/R to 0°/R in thevicinity of the blade tip.

This twist relationship may correspond to a polynomial curve, e.g. aBézier curve of order 6 or greater.

Preferably, the first plateau is equal to −18°/R, the second plateau isequal to −6°/R, the third plateau is equal to −13°/R, and the fourthplateau of the twist is equal to −8°/R at the blade tip.

Whatever the relationship for twist variation, the ninth distance mayfor example be equal to 0.35R, the tenth distance may be equal to 0.48R,the eleventh distance may be equal to 0.78R, and the twelfth distancemay be equal to 0.92R.

The twist relationship defines only variation in the twist of the bladebetween the start of the airfoil portion and the blade tip, and it doesnot define the settings of the profiles of the sections of the blade.The settings of the profiles of the sections of the blade in thevicinity of the start of the airfoil portion have no direct influence onthe aerodynamic behavior of the blade. Specifically, when in flight, thesettings of the profiles of the sections of the blade in the vicinity ofthe start of the airfoil portion and all of the profiles of the bladealong the airfoil portion depend on the adjustment of the collectivepitch and on the adjustment of the cyclic pitch of the blade. It is thusindeed the variation in twist that characterizes the aerodynamicbehavior of the blade, since the setting values of the profiles of thesections of the blades are taken into account by adjusting thecollective pitch and by adjusting the cyclic pitch of the blade.

In addition, the zones of the blade situated in the proximity of theaxis of rotation A, and in particular the zone situated between the axisof rotation A and the eighth section S8, are subjected little toaerodynamic forces during rotation of the blade. Twist in the proximityof the axis of rotation A thus has little influence on the aerodynamicbehavior of the blade. As a result, the twist may be substantiallyconstant or it may vary a little between the start of the airfoilportion and the eighth section S8 without significantly modifying thebehavior and the aerodynamic performance of the blade. By way ofexample, the variation of the twist may be less than or equal to 2°between the start of the airfoil portion and the eighth section S8.

In addition, the blade may include a dihedral beginning at the sixthsection S6 and terminating at the blade tip. This dihedral preferablyslopes downwards and serves to improve the aerodynamic performance ofthe blade in hovering flight.

The present invention also provides a rotor for a rotary wing aircraft.The rotor has at least two blades as described above. The rotor is moreparticularly intended to be a main rotor of a rotary wing aircraft forproviding it with lift and possibly with propulsion.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

The invention and its advantages appear in greater detail from thecontext of the following description of embodiments given by way ofillustration and with reference to the accompanying figures, in which:

FIGS. 1 and 2 show a blade of the invention;

FIG. 3 shows an aircraft having a rotor made up of such blades;

FIG. 4 is a graph plotting variation in the chord of the profiles of thesections of the blade;

FIG. 5 is a graph plotting variation in the sweep of the blade;

FIG. 6 is a graph plotting variation in the twist of the blade; and

FIG. 7 is a graph plotting variation in the twist gradient of the blade.

DETAILED DESCRIPTION OF THE INVENTION

Elements present in more than one of the figures are given the samereferences in each of them.

FIGS. 1 and 2 show a blade 1 extending firstly spanwise along a bladeaxis B between a blade start 2 and a blade tip 9 and secondly along atransverse axis T perpendicular to the blade axis B between a leadingedge 6 and a trailing edge 7. The blade 1 has an airfoil portion 4situated between the blade start 2 and the blade tip 9. The airfoilportion 4 is made up of a succession of airfoil profiles 15, eachsituated in a transverse plane substantially perpendicular to the bladeaxis B, each profile defining a section of the blade 1. The blade 1 alsohas a dihedral 5 at the free end of the blade 1, i.e. in the vicinity ofthe blade tip 9.

The blade 1 is for forming a rotor 11 of a rotary wing aircraft 10, asshown in FIG. 3. The rotor 11 comprises a hub 12 and five blades 1 thatare for rotating about an axis of rotation A of the hub 12. Each blade 1is connected to the hub 12 at the blade start 2.

The rotor 11 is characterized by a rotor radius R, i.e. the distancebetween the axis of rotation A and the blade tip 9 along the blade axisB. The chord c of the profiles 15 of each section of the blade 1corresponds to the maximum distance between the leading edge 6 and thetrailing edge 7 of the blade 1 in a transverse plane substantiallyperpendicular to the blade axis B. A mean chord c is defined as beingthe mean value of the chords c over the airfoil 4. The blade start 2 issituated at a fourth distance equal to 0.1R from the axis of rotation Aand the start 3 of the airfoil 4 of the blade 1 is situated at a fifthdistance equal to 0.2R from the axis of rotation A.

The blade 1 of the invention presents a combination of relationships forvariation in its sweep and in the chord of its profiles 15 firstly inorder to reduce the noise given off by each blade 1 of the rotor 11during an approach flight, and secondly in order to improve theaerodynamic performance of each blade 1 during forward flight of theaircraft 10.

Furthermore, the blade 1 may also present a combination of relationshipsfor variation in its sweep and in the chord of the profiles 15 of itssections together with a relationship for variation in its twist,firstly in order to reduce the noise given off by each blade 1 of therotor 11 during an approach flight, and secondly in order to improve theaerodynamic performance of each blade 1 both during hovering flight andforward flight of the aircraft 10.

The relationship for variation in the chord, the sweep, and the twist ofthe profiles 15 of the sections of the blade 1 are plotted respectivelyin FIGS. 4 to 6. FIG. 7 shows the twist gradient of the blade 1, i.e.the local derivative of the twist along the span of the blade 1 of therotor 11 of rotor radius R.

The relationship for variation in the chord of the profiles 15 of thesections of the blade 1 shown in FIG. 4 comprises, plotted along theabscissa axis, the ratio of the positions of the profiles 15 of thesections of the blade 1 along the span of the blade 1 relative to therotor radius R, and up the ordinate axis, the ratio of the chords c ofthe profiles 15 of the sections of the blade 1 relative to the meanchord c.

The mean chord c is defined by a radius squared r² weighting of eachprofile 15 of the sections of the blade 1 in application of thefollowing formula:

$\overset{\_}{c} = \frac{\int_{R_{0}}^{R}{{L(r)} \cdot r^{2} \cdot {dr}}}{\int_{R_{0}}^{R}{r^{2} \cdot {dr}}}$

where L(r) is the length of the local chord of a profile of the blade 1situated at a radius r from the axis of rotation A, R₀ is the radius ofthe start 3 of the airfoil 4, and R is the radius of the blade tip 9.

In this relationship for variation in the chord, the chord c of theprofile 15 of each section of the blade 1 increases between the start 3of the airfoil portion 4 and a first section S1 situated at a firstdistance from the axis of rotation A that is equal to 0.85R. Beyond thefirst section S1, the chord decreases to the blade tip 9. It can be seenthat the section c is less than the mean chord c between the start ofthe airfoil portion of the blade 1 and a fourth section S4 situated at asixth distance from the axis of rotation A equal to 0.6R. Furthermore,the chord c varies between the start 3 of the airfoil portion 4 and thefirst section S1 from 0.8c to 1.2c, which represents variation of ±20%about the mean chord c. The chord at the blade start is equal to 0.3c.

Thereafter, the chords of the profiles 15 of the sections of the blade 1are greater than the mean chord c between the fourth section S4 and afifth section S5 situated at a seventh distance from the axis ofrotation A lying in the range 0.85R to 0.95R. Finally, the chords of theprofiles 15 of the sections of the blade 1 are less than the mean chordc beyond the fifth section S5 to the blade tip 9.

In addition, the chord c decreases following a curve that issubstantially parabolic beyond a sixth section S6 situated at an eighthdistance equal to 0.95R. The end of the blade 1 thus forms a parabolictip cap 8.

The relationship for variation in the sweep of the blade as shown inFIG. 5 defines three sweeps. The ratio of the position of the profile 15of each section of the blade 1 along the blade axis B over the rotorradius R is plotted along the abscissa axis, and the sweep angle α ofeach of these profiles 15 is plotted up the ordinate axis.

Thus, the sweep is initially directed towards the front of the blade 1between the start 3 of the airfoil portion 4 and a second section S2situated at a second distance from the axis of rotation A equal to0.67R, the leading edge 6 forming a forward first sweep angle α₁ equalto 4° relative to the blade axis B. Thereafter, the sweep is directedtowards the front of the blade 1 between the second section S2 and athird section S3 situated at a third distance from the axis of rotationA equal to 0.85R, the leading edge 6 forming a forward second sweepangle α₂ equal to 8° relative to the blade axis B. Finally, the sweep isdirected towards the rear of the blade 1 between the third section S3and the blade tip 9, the leading edge 6 forming a backward third sweepangle α₃ equal to −23° relative to the blade axis B.

Each of the connections between the first, second, and third sweepangles is preferably made with a connection radius in order to avoidhaving a sharp angle at any of these connections. These connection radiimay for example be of the order of 500 millimeters (mm).

Furthermore, the blade 1 has a downwardly-directed dihedral 5 at itsfree end. This dihedral 5 begins in the vicinity of the sixth section S6and terminates at the blade tip 9. The dihedral 5 serves mainly toimprove the aerodynamic behavior of the blade 1 in hovering flight byreducing the influence of the vortex generated by the preceding blade.

In addition, a relationship for twist of the profiles 15 may be added tothe blade 1 in order to improve the aerodynamic performance of the blade1 both during hovering flight and during forward flight. Thisrelationship for twist of the blade 1 shown in FIG. 6 is a non-linearrelationship corresponding to a polynomial curve. The ratio of theposition of each profile 15 of the sections of the blade 1 along thespan over the rotor radius R is plotted along the abscissa axis, and thetwist angle θ of the profile 15 of each section of the blade 1 isplotted up the ordinate axis.

The twist gradient is shown in FIG. 7 and comprises, along the abscissaaxis, the ratio of the position of the profile 15 of each section of theblade 1 along the span of the blade 1 over the rotor radius R, and, upthe ordinate axis, the local derivative of the twist of the profile 15.

Initially, the twist angle θ varies little between the start 3 of theairfoil portion 4 and a seventh section S7 situated at a ninth distancefrom the axis of rotation A equal to 0.35R. The variation in the twistangle θ is less than 2° between the start 3 of the airfoil portion 4 andthe seventh section S7. The twist angle θ increases a little and thendecreases along the span, the twist gradient being positive in thevicinity of the start 3 of the airfoil portion 4 and decreasing tobecome negative in the vicinity of the seventh section S7.

Thereafter, the twist angle θ decreases between the seventh section S7and an eighth section S8 situated at a tenth distance from the axis ofrotation A equal to 0.48R, the twist gradient decreasing to a firstplateau equal to −18°/R in the vicinity of the eighth section S8.

Thereafter, the twist angle θ decreases less between the eighth sectionS8 and a ninth section S9 situated at an eleventh distance from the axisof rotation A equal to 0.78R, the twist gradient increases up to asecond plateau equal to −6°/R in the vicinity of the ninth section S9.In particular, the twist angle θ is equal to 0° for a profile 15 of theblade 1 situated at a distance from the axis of rotation A equal to0.65R.

The twist angle θ again decreases more between the ninth section S9 anda tenth section S10 situated at a twelfth distance from the axis ofrotation A equal to 0.92R, the twist gradient decreasing to a thirdplateau equal to −13°/R in the vicinity of the tenth section S10.

Finally, the twist angle θ decreases between the tenth section S10 andthe blade tip 9, the twist gradient increasing up to a twist gradientequal to −8°/R at the blade tip 9.

Naturally, the present invention may be subjected to numerous variationsas to its implementation. Although several implementations aredescribed, it will readily be understood that it is not conceivable toidentify exhaustively all possible embodiments. It is naturally possibleto envisage replacing any of the means described by equivalent meanswithout going beyond the ambit of the present invention.

What is claimed is:
 1. A blade for a rotor of a rotary wing aircraft,the blade being for rotating about an axis of rotation (A), the bladeextending firstly along a blade axis (B) between a blade start suitablefor being connected to a hub of the rotor and a blade tip situated at afree end of the blade, and secondly along a transverse axis (T)perpendicular to the blade axis (B) between a leading edge and atrailing edge, the blade comprising an airfoil portion situated betweenthe blade start and the blade tip, the airfoil portion being constitutedby a succession of airfoil profiles, each airfoil profile being situatedin a transverse plane substantially perpendicular to the blade axis (B)and defining a section of the blade, the blade tip being situated at adistance equal to a rotor radius R from the axis of rotation (A), amaximum distance between the leading edge and the trailing edge in thetransverse plane constituting a chord c for the airfoil profile of eachof the sections of the blade, a mean chord c being a mean value of thechord c over the airfoil portion, a forward first direction beingdefined from the trailing edge to the leading edge, and a rearwardsecond direction being defined from the leading edge to the trailingedge, the blade presenting a combination of relationships for variationin chord and in sweep, the sweep being the angle between the leadingedge and the blade axis (B), the chord increasing between the start ofthe airfoil portion and a first section S1 situated at a first distancefrom the axis of rotation (A) lying in the range 0.6R to 0.9R, the chorddecreasing beyond the first section S1; and wherein the sweep isdirected towards the front of the blade between the start of the airfoilportion and a second section S2 situated at a second distance from theaxis of rotation (A) lying between 0.5R and 0.8R, the leading edgeforming a forward first sweep angle α₁ that is strictly greater than 0°and less than 10° relative to the blade axis (B), the sweep beingdirected towards the front of the blade between the second section S2and a third section S3 situated at a third distance from the axis ofrotation (A) lying in the range 0.6R to 0.95R, the leading edge forminga forward second sweep angle α₂ lying in the range 1° to 15° relative tothe blade axis (B), the sweep being directed towards the rear of theblade between the third section S3 and the blade tip, the leading edgeforming a backward third sweep angle α₃ lying in the range −35° to −15°relative to the blade axis (B).
 2. A blade according to claim 1, whereinthe forward first sweep angle α₁ is different from the forward secondsweep angle α₂.
 3. A blade according to claim 2, wherein the forwardfirst sweep angle α₁ is strictly less than the forward second sweepangle α₂.
 4. A blade according to claim 1, wherein the forward firstsweep angle α₁, the forward second sweep angle α₂, and the backwardthird sweep angle α₃ are constant respectively between the start of theairfoil portion and the second section S2, between the second section S2and the third section S3, and between the third section S3 and the bladetip.
 5. A blade according to claim 1, wherein the forward first sweepangle α₁ is equal to 4°, the forward second sweep angle α₂ is equal to8°, and the backward third sweep angle α₃ is equal to −23°.
 6. A bladeaccording to claim 1, wherein the blade start is situated at a fourthdistance lying in the range 0.05R to 0.3R from the axis of rotation (A)and the start of the airfoil portion is situated at a fifth distancelying in the range 0.1R to 0.4R from the axis of rotation (A), the fifthdistance being greater than or equal to the fourth distance, and thechord in the vicinity of the start of the blade lying in the range 0.4cto 0.9c.
 7. A blade according to claim 1, wherein the chord varies aboutthe mean chord c by ±40% between the start of the airfoil portion andthe first section S1.
 8. A blade according to claim 1, wherein the chorddecreases in non-linear manner beyond a sixth section S6 situated at aneighth distance from the axis of rotation (A) lying in the range 0.9R to0.95R to the blade tip.
 9. A blade according to claim 8, wherein thechord decreases in parabolic manner beyond the sixth section S6.
 10. Ablade according to claim 1, wherein the blade has a dihedral in thevicinity of the blade tip.
 11. A blade according to claim 1, wherein themean chord c is defined by a radius squared r² weighting of the profileof each of the sections of the blade in application of the formula:$\overset{\_}{c} = \frac{\int_{R_{0}}^{R}{{L(r)} \cdot r^{2} \cdot {dr}}}{\int_{R_{0}}^{R}{r^{2} \cdot {dr}}}$where L(r) is the length of the local chord of a profile of the blade,the local profile being situated at a radius r from the axis of rotationA, R₀ being the radius of the start of the airfoil portion, and R beingthe radius of the blade tip.
 12. A rotor for a rotary wing aircraft, therotor having at least two blades according to claim 1.